Date of Thesis

5-1-2015

Thesis Type

Honors Thesis (Bucknell Access Only)

Degree Type

Bachelor of Science in Mechanical Engineering

First Advisor

Charles Knisely

Abstract

The vortex panel method, using 253 panels, was used to compute the circulation and pressure distribution around an NACA 2412 airfoil with a 10 in. chord. A similar calculation was carried out with two additional bodies representing wind tunnel walls bounding the flow, equally disposed above and below the airfoil. Calculations for both flow geometries were carried out over ranges in freestream velocity from 0.1 to 100 m/s, resulting in a Reynolds number range based on the airfoil chord of 1,675 to 1,675,704. The angle of attack was varied from -15° to 15°. Boundary layer analysis was carried out over the airfoil surface using Thwaites' method for laminar flow, Michel's method for transition, and Head's method for turbulent flow. The computed viscous lift and drag were added to the pressure lift and drag to determine the total lift and drag on the airfoil as a function of the angle of attack, for both the isolated airfoil in unbounded flow and the airfoil between simulated walls with a bounded flow. The calculated lift and drag coefficient data for the bounded flow showed effects of wind tunnel blockage, as anticipated. Two published blockage correction methods, one from Maskell and the other from Selig, were applied to the bounded flow values; the corrected data was, for the most part, in substantial agreement with the unbounded calculated values. With corrections from Maskell's method, the corrected pressure and drag to values approached the corresponding unbounded data, but were not exactly the same. The corrected lift values form Maskell's method had negligible difference relative to the unbounded data. Using Selig's method, the corrected values matched the unbounded data for positive angles of attack. For negative angles of attack, however, the corrected values were further from the unbounded data than the original bounded data.

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